[0001] The present invention relates to aerofoils and rotor or stator assemblies of such
aerofoils, typically for use in turbomachinery.
[0002] This invention is particularly concerned with improving the control of the fluid
flow past one or more rows of such aerofoil members, which may be fixed vanes or may
be blades rotating about the central axis of the duct. All aerofoils, whether in compressors
or in turbines, use surface curvature to change the static pressure of the flow and
thus provide lift. All such blade rows suffer from secondary flows that arise at the
end walls of the blades increasing loss generation.
[0003] The asymmetry between the pressure and suction surface of an aerofoil (i.e. the aerofoil
camber) produces lift. Blades and vanes in early compressors were designed with circular
arc camber lines over the whole span of the blade such that the camber is evenly distributed
along the entire chord. However the location of peak aerodynamic loading along the
chord length caused by such a camber distribution has been found to result in high
losses and a sub-optimal operating range for the flow through the compressor. Developments
in subsonic aerofoil design have seen the camber distribution being pushed forward
from the initial circular arc designs. Thus greater camber is provided in the leading
portion of the blade than in the trailing portion.
[0004] Friction between the working fluid (typically air) and a solid surface, such as the
aerofoil surface and/or the wall of a fluid passageway, creates a boundary layer of
slower moving air. The faster air flowing between the aerofoils, outside the boundary
layer, is in equilibrium with the pressure gradient caused by the aerofoils and is
turned to the desired exit flow angle. However the boundary layer air passing between
the aerofoils is more strongly influenced by the pressure gradient between the opposing
surfaces of adjacent blades, such that it is turned to a greater degree than the primary
air flow and rolls up into vortices, creating secondary flows that result in increased
aerodynamic losses. (See, for example,
Takeishi et al. (1989, "An Experimental Study of the Heat Transfer and Film Cooling
on Low Aspect Ratio Turbine Nozzles", ASME Paper 89-GT-187).
[0005] In compressors, the problems associated with these secondary flows are exacerbated
because generally compressor blade rows diffuse the flow. Because of this, typically
the over-turned boundary layer may not simply roll up into a wake from the trailing
edge of the blade, but additionally a region of separated flow can form in the corner
between the aerofoil suction surface and the end wall. In parts of this separated
region, the airflow may be reversed.
[0006] This corner separation is a source of significant losses, typically larger than the
losses arising from "standard" secondary flows (such as in turbines). Corner separation
may also cause significant blockage of the effective flow area. That is to say a portion
of the available area of the flow passage is consumed by the vortices, thereby restricting
the area available for the primary flow and reducing the mass flow delivered by the
compressor.
[0007] As the pressure rise demanded of a compressor increases, it moves up its "characteristic".
The aerofoils experience positive incidence and increased aerodynamic loading. Under
these conditions the corner separation rapidly increases and may initiate stalling
of the blade row. This behaviour may ultimately define the surge margin of the machine.
[0008] The front stages of compressors experience the most extreme off-design conditions
(i.e. increased airflow incidence angles). If the pressure ratio of the compressor
is large enough, variable stator vanes (VSVs) are required to mitigate the extreme
incidence onto front stage rotors. The lower in speed the machine operates relative
to design speed, the more the variable vanes are closed. However employing VSVs has
drawbacks. Apart from the extra cost and weight of VSV's, additional complexity is
introduced by the need to implement suitable control systems. Also re-distribution
of mass flow impairs the effectiveness of variable vanes. Typically this leads to
one set of endwalls being preferentially over-loaded aerodynamically.
[0009] Achieving surge margin requirements is of considerable importance for gas turbine
engine compressors, and represents a particular problem at part speed where surge
can initiate from the compressor endwall regions. One way of meeting challenging surge
margin requirements at part speed is by modifying the VSV schedule. However this solution
significantly reduces part speed efficiency and can lead to increased rotor noise
(so-called "pseudo stall"). The surge margin of the compressor also becomes sensitive
to VSV malschedule, which is not ideal.
[0010] Another known means of controlling corner separation is by three-dimensional leaning
of the aerofoils. The application of "sweep" and "dihedral" leans (parallel and normal
to the aerofoil chord line, respectively) can reduce the separation and flow reversal
seen near to the end walls, and this improves the design point efficiency, compressor
pressure ratio and surge margin of compressors.
[0011] However, there are drawbacks. The scope for leaning the aerofoils (particularly those
in rotating blade rows) may be limited by mechanical constraints, especially stress.
Even in stationary components (stators and vanes) where stress is not so important,
manufacturing constraints may limit the three-dimensional shaping. Also, if the aerodynamic
loading of the compressor and its blade rows is high enough, even the optimum leant
aerofoil shape may not be enough to eliminate the corner separation.
[0012] An alternative means of controlling the flow regime are disclosed in patents
US 6,283,713 and
US 7,354,243 (both in the name of Rolls-Royce plc), which disclose a sinusoidally varying surface
profile applied to the end wall of a turbomachinery blade row to create variations
in the axial or circumferential direction. The static pressure distribution of the
flow on the end wall is thus modified in order to control the secondary flow.
[0013] It is an aim of the present invention to provide an alternative means for mitigating
the effects of aerofoil boundary layer separation in an axial flow machine. According
to a first aspect of the invention there is provided an aerofoil for an axial flow
machine, the aerofoil having first and second ends and opposing pressure and suction
surfaces extending between said first and second ends in a span-wise direction, wherein
in a first portion of the span of the aerofoil towards the first end the aerofoil
has a location of greatest camber which is closer to a trailing edge than a leading
edge of the aerofoil and wherein in a further portion of the aerofoil span the camber
is either uniform between the leading and trailing edges or else a location of greatest
camber is closer to the leading edge than the trailing edge.
[0014] According to a second aspect of the invention there is provided an aerofoil for an
axial flow machine, the aerofoil having first and second ends and opposing pressure
and suction surfaces extending between said first and second ends in a span direction,
wherein in a first portion of the span of the aerofoil towards the first end the pressure
and/or suction surface turns to a greater extent towards a trailing edge than towards
a leading edge of the aerofoil and wherein in a further portion of the aerofoil the
pressure and/or suction surface turns uniformly between the leading and trailing edges
or else turns to a greater extent towards a leading edge.
[0015] The turning of the surfaces may be defined with respect to turning within a cross-sectional
plane. For example a greater curvature (or smaller radius of curvature) of the pressure
and/or suction surface in a cross-sectional plane may equate to a greater degree of
turning. A smaller curvature (or greater radius of curvature) of the pressure and/or
suction surface in a cross-sectional plane may equate to a smaller degree of turning.
[0016] In one embodiment the first portion is immediately adjacent the first end.
[0017] The first portion may encompass or extend between 1 and 50% of the span of the aerofoil.
The first portion may encompass or extend between 1 and 30% of the span of the aerofoil.
The first portion may encompass or extend between less than or approximately equal
to 25% of the span of the aerofoil, such as between 5% and 20% of the aerofoil span
or else between 10% and 20% of the span of the aerofoil.
[0018] The invention may otherwise be defined as the provision of an axial flow machine
aerofoil having a rearwardly loaded camber distribution near the first end of the
aerofoil. The provision of such an aerofoil, or array thereof, is beneficial in that
the corner separation for a given level of aerodynamic loading can be reduced and/or
the stable operating range of fixed rotor/stator endwall regions can be increased.
This is contrary to the common understanding that it is beneficial to have a forwardly
loaded camber distribution over the span of the blade in order to improve efficiency
by minimizing losses associated with the primary flow (2D profile).
[0019] The further portion may be adjacent the first portion. The further portion of the
aerofoil span may comprise the remainder of the span (i.e. the remainder of the blade
between the first portion and the second end). For example the further portion may
comprise between 50% and 90% of the aerofoil span, or, more typically, between 70%
and 80% thereof.
[0020] The further portion may comprise a second portion which is adjacent the first portion.
The second portion may have a uniform camber between the leading and trailing edges,
or else a camber which is greatest a central location between the leading and trailing
edges of the aerofoil. In the second portion the pressure and/or suction surfaces
may turn uniformly between the leading and trailing edges or else may turn to a greatest
extent (or at a greatest rate) at a central location between the leading and trailing
edge.
[0021] The further portion may consist of or comprise of a third portion. The third portion
may be spaced from the first portion by the second portion. A location of greatest
camber in the third portion may be closer to the leading edge than the trailing edge.
The pressure and/or suction surface in the third portion may turn to a greater extent
towards the leading edge than towards the trailing edge.
[0022] The aerofoil may be arranged to depend from an end wall at the first end such that
the aerofoil is upstanding therefrom in use. The aerofoil may be arranged to span
an annular flow passage such that an end wall is provided at both the first and second
ends thereof in use.
[0023] The aerofoil is typically arranged for subsonic flow speeds in use.
[0024] The first end may comprise a radially inner end of the aerofoil with respect to an
axis of rotation of the axial flow machine. The location of greatest camber may tend
towards the leading edge of the aerofoil with distance from the first end, for example
within the first portion and/or further portion. Over the further portion of the aerofoil
(such as a portion towards the second end or else a mid-span portion of the aerofoil),
the location of greatest camber may be substantially constant.
[0025] Alternatively, over a portion of the aerofoil towards the second end, the location
of greatest camber may again be closer to a trailing edge than a leading edge of the
aerofoil (e.g. akin to the first portion towards the first end). Therefore a first
portion may be provided at both of the first and second ends of the aerofoil, with
the further portion located there-between.
[0026] That is to say the invention can, for example, be applied towards one or both ends
of an aerofoil.
[0027] The present invention is particularly useful in that it can also be applied in conjunction
with the further means for means reducing secondary flows as discussed above in the
introduction.
[0028] According to a third aspect of the invention, there may be provided a rotor or stator
aerofoil assembly for an axial flow machine comprising a plurality of aerofoils according
to the first or second aspect arranged in an annular array about an axis. A row of
rotor blades and/or stator vanes may be provided, wherein each blade or vane is according
to the first or second aspect.
[0029] Each aerofoil in the array may have a pressure surface which faces a suction surface
of an adjacent aerofoil.
[0030] The invention may reduce the three-dimensional flow (or separated boundary layer)
interaction between the flow over the suction surface of one aerofoil and the flow
over the pressure surface of an adjacent aerofoil. The present invention may improve
the aerodynamic stability of an axial flow compressor or turbine.
[0031] The assembly may comprise end walls defining an annular flow passage, wherein the
aerofoils extend across the flow passage between the end walls in a generally radial
direction with respect to the axis. The axis is typically common to the rotor and
stator and may comprise a longitudinal axis or axis of rotation for the assembly.
[0032] According to a fourth aspect of the invention there is provided a turbomachinery
compressor or turbine, such as a gas turbine engine, comprising one or more blades
according to the first or second aspect or else an assembly according to the third
aspect. There may be provided a gas turbine engine having such a compressor or turbine.
[0033] Any of the preferable features defined in relation to any one aspect of the invention
may be applied to any further aspect.
[0034] Practicable embodiments of the invention are described in further detail below by
way of example only with reference to the accompanying drawings, of which:
Figure 1 shows a three-dimensional view of the flow over adjacent aerofoils according
to the prior art;
Figure 2 shows a side view of an aerofoil according to the invention;
Figure 3 shows a cross section through a first portion of adjacent aerofoils according
to one example of the invention;
Figure 4 shows a cross section through a further portion of an aerofoil of Figure
3;
Figure 5 shows a cross section through another portion of an aerofoil of Figure 3;
Figure 6a-c show plots of camber distribution for the cross sections of Figures 3,
4 and 5 respectively;
Figure 7 shows a three-dimensional view of computational fluid dynamics results of
the flow over adjacent aerofoils for a uniform camber profile;
Figure 8 shows a three-dimensional view of computational fluid dynamics results of
the flow over adjacent aerofoils for one example of an aerofoil assembly according
to the invention;
Figure 9 shows a plot of lift distribution for a first portion of a compressor stator
according to one example of the invention;
Figure 10 shows plots of loss coefficient against inlet whirl angle for the stator
of Figure 9; and,
Figure 11 shows a longitudinal half-section of a gas turbine engine according to the
invention.
[0035] Figure 1 shows the results from a Computational Fluid Dynamics (CFD) simulation of
the flow in a simple linear cascade of compressor blade rows. Two adjacent aerofoils
(10a, 10b) are shown, each having a suction surface (respectively 12a, 12b) and a
pressure surface (14a, 14b). A platform (16a, 16b) is associated with each aerofoil
(10a, 10b), the platforms (16a, 16b) being radially inward of the aerofoils and together
defining an end wall 17. A flow passage (18) is defined by the surfaces 14a, 17 and
12b and by a shroud surface (not shown) radially outward of the aerofoils (10a, 10b).
The model is viewed from the rear (from downstream looking upstream). Streamlines
show the flow patterns on the end wall (17) and on the aerofoil suction surface (12b)
for the lower half of the passage). The incoming end wall boundary layer (20) approaches
the leading edge at the design incidence.
[0036] The over-turning of the fluid at the bottom of the inlet boundary layer is visible
in the streamlines (22). This causes the airflow to impinge onto the suction surface
(12b) of the aerofoil. On the aerofoil, near mid-height, the flow lines are two-dimensional
(24). However, nearer the end wall (17) the streamlines on the aerofoil are highly
three-dimensional and clearly indicate the region of separated flow (26) (where some
reverse flow can be seen).
[0037] Turning now to Figure 2, there is shown an aerofoil in the form of a stator vane
28, to which the principles of the present invention are applied. A series of twenty-one
sections 30 are equally spaced locations along the span 32 of the vane so as to define
the overall geometry of the vane. Each section 30 typically lies along a streamline.
However such sections may be referred to herein as cross-sections.
[0038] A radially inner portion of the vane A, which may be referred to as a hub region,
extends from the hub end 34 of the vane approximately 25% of the way along the span
of the vane. A mid-span region B occupies between 25%-75% of the span of the vane
and represents a "free stream" region in which the flow over the vane is substantially
two-dimensional (i.e. less affected by the end walls). A casing region C of the vane
extends over the final quarter of the vane span (i.e. from 75%-100% of the span) in
the vicinity of the radially outer end 36 of the vane.
[0039] The leading 38 and trailing 40 edges of the vane extend the full length of the span
32 and represent respective front and aft extremities of the vane in the direction
of flow thereover in use.
[0040] Turning now to Figure 3 there is shown a section 30A taken at the hub end 34 of a
pair of adjacent vanes 28 within a compressor. Each section 30A is juxtaposed against
a section 42 through a vane having a circular arc camber line, indicative of uniform
turning of the vane between its leading and trailing edges. In Figure 3 it can be
seen that the vane 28 at its hub end turns less in the leading/front half of the vane.
Accordingly the rate of turning of the vane increases to a maximal value in the rear/trailing
half of the vane. Thus the angle through which the section 30A of the vane turns (i.e.
its curvature) in the front half is lower than in the rear half.
[0041] This aerofoil geometry can be further explained with reference to Figure 6a, which
plots at line 44 the camber distribution (i.e. the change in camber against distance
along the chord length) for section 30A. That camber distribution 44 is compared to
a camber distribution for a circular arc (i.e. uniform camber) at line 46. It can
be seen that the gradient of the line 44 is less than that of the line 46 towards
the front half of the aerofoil (i.e. below a value of 0.5 on the Y-axis), being indicative
of a lower camber, whereas the gradient of line 44 is greater than that of 46 in the
rear half of the aerofoil. In this example only 20-25% of the camber (or blade turning)
occurs in the front half of the blade in the cross-sectional plane. Approximately
half of the camber is achieved at substantially 75% of the chord length.
[0042] Turning now to Figure 4, there is shown an aerofoil section 48 taken at the fourth
of the sections 30 along the span of the vane 28. Figure 6b shows the camber profile
for section 48 at line 50. For comparison, the camber distributions for a circular
arc 46 and also a conventional forward-loaded vane camber profile 52 are shown. Here
it can be seen that the camber profile 50 of the vane 28 is generally similar to that
of the circular arc profile 46. However the profile 50 is slightly forwardly-loaded,
such that by section 48 (i.e. at some point between the third and fourth sections
along the vane span), the vane profile has passed through a uniform camber profile.
Thus the location of uniform camber corresponds to a location of between 15% and 20%
of the distance along the blade span in this embodiment.
[0043] The region where transition from rearward to forward-loaded camber style occurs on
an aerofoil requires care. The ideal aerofoil would have healthy suction surface boundary
layers across the entire span of the aerofoil akin to forward loaded aerofoils, with
endwall boundary layer overturning occurring as late as possible in the passage. In
practice, near the endwall, the suction surface boundary layer health has to be compromised
to allow for the rearward loaded camber style. The key to the transition point is
to minimise disruption to the suction surface boundary layers in the 2D region but
yet still have sufficient influence over the endwall wall flow.
[0044] This transition point location may be varied for different blade or vane geometries
but a suitable location is generally considered to lie between 10% and 30%, and, more
typically, around 20-25% of span inboard from the endwalls. At such a location it
has been found that the benefits of the delayed roll up of endwall boundary layer
flow onto the suction surface outweighs the detriment of the resulting less healthy
suction surface boundary layers. Transitioning too close to the endwalls would result
in little gain, whereas transitioning too far inboard of the endwall could see a rise
in loss and potentially a reduction in stable operating range due to the disruption
caused to the boundary layers in the 2D region.
[0045] In Figure 5 there is shown a section 54 of the vane taken at the seventh of the aerofoil
sections 30 shown in Figure 2. Figure 6c shows the camber plot 56 for this section,
which is substantially equal to a forward-loaded aerofoil of this type according to
the prior art (i.e. closer to the line 52 shown in Figure 6b). Further discussion
of this camber distribution 52 will be omitted for brevity. However it is noted that
any of the details of the camber distributions shown in any of Figures 6a-c may contribute
to the definition of the invention.
[0046] A more-conventional forward-loaded camber distribution, as shown in Figures 5 and
6c is adopted over the remainder of the span of the aerofoil, i.e. for 20-75% of the
span in this embodiment towards the radially outer end of the vane.
[0047] However in alternative embodiments it is possible to apply the rearward-loaded camber
distribution at each end of the aerofoil such that a forward-loaded camber distribution
at the central region is bounded at each end by a rearward-loaded camber distribution
region, typically in the vicinity of the end walls. Each of those regions may be as
described above. However it is not essential that those regions are identical. For
example, if the boundary layer effects are reduced at the radially outer endwall,
then the length of the rearward-loaded camber region at that end may be reduced. Furthermore
the degree of rearward loading itself may be reduced. Alternatively a radially outer
end of the aerofoil may not require a rearwardly-loaded camber distribution but may
instead have only a circular arc camber profile or else a less-forwardly-loaded camber
profile than the mid region of the aerofoil.
[0048] Turning now to Figures 7 and 8, there are shown 3D Computational Fluid Dynamics (CFD)
flow visualisation for adjacent stator vanes of an intermediate pressure compressor
of a gas turbine engine. Although the flow profile is different to that of Figure
1 corresponding sources of aerodynamic losses can be seen. Figure 7 shows a flow visualization
for a pair of conventional vanes 58. The streamlines 60 on the aerofoil surface and
on the end wall 62 show that the endwall boundary layer rolls up to form a vortex
64 which passes downstream, causing a wake region (shaded area 66) between the pressure
and suction surfaces of the adjacent vanes.
[0049] Also shown in Figures 7 and 8 are corresponding contour plots of total pressure at
an outlet plane of the stator row. The total pressure in the free stream between the
two vanes at location 68 is approximately 200 kPa, whereas at its lowest value, in
the region of the point 70, the boundary layer wake causes a drop in pressure to approximately
170 kPa.
[0050] Figure 8 shows the same CFD visualization for the vane described above in relation
to Figures 2-6c. Here it can be seen that the modified vane has a smaller wake region
72, indicative of reduced corner separation. In this regard the wake region may be
substantially halved in size. The minimum total pressure at the exit plane at location
74 of Fig. 8 is greater than that of Fig. 7 and is approximately equal to the pressure
denoted by contour 76 in Fig. 7. Thus it can be seen that the area of separated flow
in Fig. 8 is reduced in size. The area or volume by which the wake flow can be reduced
may be in the region of 50% or more.
[0051] Furthermore the location 78 of the onset of separated flow is moved downstream along
the suction surface of Fig. 8, when compared to location 64 in Fig. 7. All of the
above described features of Fig. 8 will be understood by the skilled person to be
aerodynamic improvements. Further quantification of the benefits of such revised vane
geometry is given in Figures 9 and 10.
[0052] Figure 9 shows a comparison of the lift distributions taken from 3D CFD studies for
a stator row 4 of a gas turbine engine intermediate pressure compressor in the vicinity
of the hub end wall (e.g. at approximately 5% of the vane span). It is clear that
the static pressure field near the hub has been significantly influenced by the local
change in camber distribution. In particular the mach number of the flow of the conventional
vane plot 80 peaks between 5-10% (approximately 8%) higher than the peak of the plot
82 for the vane proposed as an example of the invention. The peaks are also spaced
along the axial chord lengths of the aerofoils such that the plot 82 peaks downstream
of plot 80 at approximately 19% of the chord length. The plot 82 shows greater lift
generated in the rear half of the vane than the plot 80.
[0053] Figure 10 shows the variation of aerodynamic loss (i.e. the Streamline loss coefficient)
with change in inlet/whirl angle near the hub endwall. The loss coefficient in this
example is defined as:
where
Pt is Total Pressure and
Ps is Static Pressure.
[0054] In this plot the mean values are taken in the 0-25% mass flow region from the end
wall. As the compressor operates off-design for a given spool speed, the incidence
onto the stator increases and loss rises. The design point is shown at 83. The conventional
vane design plot is given at 84 whereas the modified design as described above is
given at 86. The new design shows that the rate of increase in loss due to increasing
incidence angle is reduced for this particular vane. It will be understood that as
the loss coefficient tends to a vertical line, the flow becomes unstable and the associated
stability limits 85 of each design have been marked. In this example, towards the
right hand end of the plots 84 and 86, it can be seen approximately 1 degree extra
stable operating range is achieved. If one degree of range is added to the endwall
region of each row, this has the potential of increasing surge margin by approximately
5%.
[0055] The present invention is also beneficial in that it provides a further means for
improving the stability of compressors, which can be used in conjunction with the
any or any combination of existing techniques. That is to say the present invention
may be applied to a blade or vane that is either swept and/or has a dihedral lean
as described above.
[0056] The present invention can additionally or alternatively be used in conjunction with
the stator hub terracing techniques disclosed in patents
US 6,283,713 and
US 7,354,243.
[0057] The present invention, whether used alone or in combination with the other techniques
described above has the potential to increase range sufficiently in compressors to
allow for a conventional VSV schedule to be adopted instead of a less-ideal, modified
schedule. For example, by using the present invention, the VSV for a downstream (e.g.
second) compressor stage may be controlled to close by less than that of the VSV of
an upstream (e.g. first) compressor stage. For example VSV2 would preferably close
by about half the amount of VSV1.
[0058] With reference to Figure 11, a ducted fan gas turbine engine is shown, in which the
present invention may be applied. The engine, generally indicated at 110, has a principal
and rotational axis 111. The engine 110 comprises, in axial flow series, an air intake
112, a propulsive fan 113, an intermediate pressure compressor 114, a high-pressure
compressor 115, combustion equipment 116, a high-pressure turbine 117, and intermediate
pressure turbine 118, a low-pressure turbine 119 and a core engine exhaust nozzle
120. A nacelle 121 generally surrounds the engine 110 and defines the intake 112,
a bypass duct 122 and a bypass exhaust nozzle 123.
[0059] The gas turbine engine 110 works in a conventional manner so that air entering the
intake 112 is accelerated by the fan 113 to produce two air flows: a first air flow
into the intermediate pressure compressor 114 and a second air flow which passes through
a bypass duct 122 to provide propulsive thrust. The intermediate pressure compressor
114 compresses the air flow directed into it before delivering that air to the high
pressure compressor 115 where further compression takes place. The intermediate 114
and high 115 pressure compressors comprise an annular passageway defined by a hub
end wall of the compressor rotor and a, radially outer casing end wall. The rotor
blades depend outwardly across the passage from the rotor end wall and the stator
vanes typically depend radially inwardly from the casing.
[0060] The compressed air exhausted from the high-pressure compressor 115 is directed into
the combustion equipment 116 where it is mixed with fuel and the mixture combusted.
The resultant hot combustion products then expand through, and thereby drive the high,
intermediate and low-pressure turbines 117, 118, 119 before being exhausted through
the nozzle 120 to provide additional propulsive thrust. The high, intermediate and
low-pressure turbines 117, 118, 119 respectively drive the high and intermediate pressure
compressors 115, 114 and the fan 113 by suitable interconnecting shafts.
[0061] Alternative gas turbine engine arrangements may comprise a two, as opposed to three,
shaft arrangement and/or may provide for different bypass ratios. Other configurations
known to the skilled person include open rotor designs, such as turboprop engines,
or else turbojets, in which the bypass duct is removed such that all air flow passes
through the core engine. The various available gas turbine engine configurations are
typically adapted to suit an intended operation which may include aerospace, marine,
power generation amongst other propulsion or industrial pumping applications.
[0062] This invention is best suited to high aspect ratio blading where it is possible to
decouple the endwall pressure field from the mid-span. Accordingly the above described
modifications to aerofoil profile may be applied to rotor blades or stator vanes in
either of the IP or HP (114, 115) compressors. The present invention may also be applied
to one or more turbine stages, such as the HP or IP turbine blades or vanes.
1. An aerofoil for an axial flow machine, the aerofoil having:
first (34) and second (36) ends and
opposing pressure and suction surfaces extending between said first and second ends
in a span direction,
wherein in a first portion of the span of the aerofoil towards the first end the aerofoil
has a location of greatest camber which is closer to a trailing edge (40) than a leading
edge (38) of the aerofoil
and wherein in a further portion of the aerofoil span the camber is either uniform
between the leading and trailing edges or else a location of greatest camber is closer
to the leading edge than the trailing edge.
2. An aerofoil according to Claim 1, wherein the first portion is at or immediately adjacent
the first end.
3. An aerofoil according to Claim 1 or 2, wherein the first portion comprises between
5% and 30% of the span of the aerofoil.
4. An aerofoil according to any preceding claim, wherein the further portion is adjacent
the first portion and comprises a remainder of the aerofoil span.
5. An aerofoil according to any preceding claim, wherein the further portion comprises
a second portion which is adjacent the first portion, the second portion having a
uniform camber between the leading and trailing edges, or else a camber which is greatest
at a central location between the leading and trailing edges.
6. An aerofoil according to Claim 5, wherein the further portion comprises a third portion,
the third portion being spaced from the first portion by the second portion, wherein
the location of greatest camber in the third portion is closer to the leading edge
than the trailing edge.
7. An aerofoil according to Claim 6, wherein the location of greatest camber is substantially
constant over the third portion of the aerofoil.
8. An aerofoil according to any one of Claims 1 to 6, wherein the location of greatest
camber within the further portion tends towards the trailing edge of the blade in
the vicinity of the second end of the aerofoil.
9. An aerofoil according to any preceding claim, wherein the location of greatest camber
within the first portion tends towards the leading edge of the aerofoil with distance
from the first end.
10. An aerofoil according to any preceding claim, wherein the opposing pressure and suction
surfaces extend from an end wall (72) at the first end.
11. An aerofoil according to Claim 10 arranged to span an annular flow passage defined
by the end wall at the first and an opposing end wall at the second end.
12. An aerofoil according to any preceding claim wherein the first end comprises a radially
inner end or hub end with respect to an axis of rotation of the axial flow machine.
13. An aerofoil according to any preceding claim wherein the aerofoil is swept and/or
leant in a dihedral direction between its first and second ends.
14. A rotor or stator aerofoil assembly for an axial flow machine comprising a plurality
of aerofoils according to any one of claims 1 to 13 arranged in an annular array about
an axis.
15. A gas turbine engine comprising a compressor having a plurality of rows of rotor blades
in axial flow series and a plurality of rows of stator vanes therebetween, wherein
the at least one row of blades and/or vanes comprises aerofoils in accordance with
any one of Claims 1 to 13.